1. Field of the Invention
The present invention relates generally to spacecraft propulsion systems and, more particularly, to ion thrusters.
2. Description of the Related Art
On-board propulsion systems are used to realize a variety of spacecraft maneuvers. In satellites, for example, these maneuvers include the processes of orbit raising (e.g., raising from a low Earth orbit to a geostationary orbit), stationkeeping (e.g., correcting the inclination, drift and eccentricity of a satellite's orbit) and attitude control (e.g., correcting attitude errors about a satellite's roll, pitch and yaw axes).
The force exerted on a spacecraft by a propulsion system's thruster is expressed in equation (1) ##EQU1## as the product of the thruster's mass flow rate and the thruster's exhaust velocity. Equation (1) also shows that mass flow rate can be replaced by the ratio of weight flow rate to the acceleration of gravity and that the ratio of exhaust velocity to the acceleration of gravity can be represented by specific impulse I.sub.sp which is a thruster figure of merit. Equation (1) can be rewritten as equation (2) ##EQU2## to show that specific impulse is the ratio of thrust to weight flow rate.
When a thruster is used to effect a spacecraft maneuver, a velocity increase .DELTA.V of the spacecraft is gained with a loss in mass of stored fuel. Thus, there will be a differential between the spacecraft's initial mass M.sub.i (prior to the maneuver) and the spacecraft's final mass M.sub.f (after the maneuver). This mass differential is a function of the thruster's specific impulse I.sub.sp as expressed by the "rocket equation" of ##EQU3## in which .DELTA.V has units of meters/second, I.sub.sp has units of seconds and a constant g is the acceleration of gravity in meter/second.sup.2. Equation (3) states that fuel loss causes a spacecraft's final mass M.sub.f to exponentially decrease with increased .DELTA.V and that this decrease can be exponentially offset by an increase in specific impulse I.sub.sp.
Specific impulse is an important measure of a thruster's fuel efficiency. Typical specific impulses are 230 seconds for monopropellant (e.g., hydrazine) thrusters, 290 seconds for solid propellant thrusters, 445 seconds for bipropellant (e.g., liquid hydrogen and liquid oxygen) thrusters and 500 seconds for electric arc jet thrusters. In contrast, ion thrusters have been developed with specific impulses in excess of 2500 seconds.
The high specific impulse of ion thrusters makes them an attractive thruster for spacecraft maneuvers. Their high fuel efficiency can facilitate a reduction of initial satellite mass, an increased payload and a longer on-orbit lifetime. Reduction of initial mass lowers the spacecraft's initial launch cost and increased payload and longer lifetime increase the revenue that is generated by the spacecraft.
The high specific impulse of ion thrusters is accompanied by thrust levels (e.g., .about.18 millinewtons in a thruster with a diameter of .about.13 centimeters) which are typically less than those of more conventional thrusters. For most spacecraft maneuvers, however, these lower thrust levels are easily accommodated by increased thruster firing times. In fact, the lower thrust levels of ion thrusters can improve satellite positioning accuracy because they facilitate frequent firings. The higher thrust levels of other thruster types necessitate less frequent firings with consequent decrease in positioning resolution.
However, their longer firing times increase the lifetime requirements of ion thrusters. In a typical satellite lifetime, for example, one of the most demanding satellite maneuvers (north-south stationkeeping) requires an ion thruster lifetime in excess of 10,000 hours. Orbit raising maneuvers can further increase this requirement. Lifetimes of these magnitudes have been difficult to obtain because of cross-over ion erosion in the ion-optics system of conventional ion thrusters. The sources of this erosion are theorized to occur as shown in FIGS. 1A-1C.
These figures illustrate the formation of exemplary ion beamlets by an array of aperture sets in a typical ion-optics system. FIG. 1A shows an aperture set 20 which includes a screen aperture 21 in a screen grid 22, an accelerator aperture 23 in an accelerator grid 24 and a decelerator aperture 25 in a decelerator grid 26. Similarly, FIG. 1B shows an aperture set 30 of a screen aperture 31, an accelerator aperture 33 and a decelerator aperture 35 and FIG. 1C shows an aperture set 40 of a screen aperture 41, an accelerator aperture 43 and a decelerator aperture 45. The aperture sets 20, 30 and 40 are positioned progressively further from the center of the aperture set array.
The screen apertures 21, 31 and 41 facilitate the flow of ion beamlets 46, 48 and 50 from a plasma sheath 52 of an ion source (each line in the beamlets indicates a different ion trajectory). Each accelerator aperture is positioned relative to its respective screen aperture so that an accelerator voltage on the accelerator grid 24 attracts the accelerator aperture's respective ion beamlet and accelerates it through the accelerator aperture. Each decelerator aperture is positioned relative to its respective screen aperture so that a decelerator voltage on the decelerator grid 26 exerts a collimating force on the decelerator aperture's respective ion beamlet.
The plasma density of the plasma source typically decreases towards the perimeter of the aperture set array and, therefore, the plasma sheath 52 extends further from the screen grid 22 and initiates increasingly angled ion trajectories. This radial decrease of plasma density also causes a corresponding decrease in the ion densities of the beamlets and, thus, a decrease in their positive space charges which tend to radially expand the beamlets.
Because of the cumulative effects of these variations, beamlet 46 passes through its aperture set, beamlet 48 begins to exhibit some crossover in its ion trajectories and several ion trajectories of beamlet 50 terminate on the decelerator grid 26. Ions on these latter trajectories sputter atoms from the decelerator grid.
In tests, this sputtering has been observed to erode a decelerator grid of an ion-optics system in operational test times (e.g., .about.500 hours) far less than the lifetime requirements cited above. Accordingly, cross-over ion erosion has prevented the realization of the lifetime requirements cited above. In addition, sputtered atoms from the decelerator grid may be deposited on sensitive spacecraft surfaces, e.g., solar cells.